For the various mission instruments of a satellite, such as the telecommunication systems, to be used, the position and the orientation of the satellite in orbit need to be controlled. In addition to the earth's gravity, a satellite is subject to a number of lower amplitude forces which progressively modify its position and its orientation. The gravity of the sun and of the moon, the deformation of the earth at the equator, or even the solar radiative pressure generate drifts which have to be corrected. Control systems are implemented to maintain, on the one hand, the orientation of the satellite relative to the earth, i.e. attitude control, and on the other hand its position in orbit relative to a desired ideal position, i.e. orbit control.
For example, in the case of geostationary satellites, orbit control seeks to limit the inclination relative to the equatorial plane, to limit the eccentricity of the orbit, and to limit the drift of the longitudinal position of the satellite relative to the earth. For this, thrusters are positioned at various points on the satellite to correct the trajectory at fairly close intervals by applying a force to the satellite. These station keeping operations require the capability to have a sufficient reserve of propellant throughout the life of the satellite. To limit the cost and the weight of the orbit control devices, a novel architecture of the propulsion systems is envisaged.
A satellite is placed in orbit by a combination of a space launch vehicle and its own propulsion systems. The launch vehicle transports and releases the satellite on a first so-called transfer earth orbit, the perigee of which is generally low; once on this first orbit, a propulsion system of the satellite takes over to transport the satellite to its final orbit. Generally, this transfer is performed by means of a main satellite thruster PSP consuming a chemical propellant of ergol or propergol type, delivering a high-power thrust in order to rapidly reach the final orbit.
Once on station, a plurality of lower power secondary thrusters ensure that the satellite is held in position on the orbit. For this, ergol chemical thrusters or electrical thrusters can be used. In an electrical thruster, of plasma thruster or ion thruster type, xenon atoms are ionized by collision with electrons, creating xenon ions. The thrust is generated when the charged xenon ions are accelerated out of the thruster by an electromagnetic field. Although costly and initially very heavy, the effectiveness of the thruster, or its capacity to generate force by mass ejection, also called specific impulse, is substantially greater than that of the chemical thrusters.
In the known systems, chemical thrusters and electrical thrusters are positioned in a number of positions on the structure of the satellite in order to meet all the needs of the mission, from transporting from the transfer orbit to maintaining in orbit throughout the life of the satellite. The propulsion systems thus implemented have the drawback of a high cost and the high weight of the various thrusters and of the propellant. These drawbacks limit the payload installation capacity of the satellite.
According to the known state of the art, an orbit control system seeks to control the position of the satellite through six orbital parameters. FIG. 1 represents a geostationary satellite 10 in orbit 11 around the earth 12. The orbit 11 is inclined by an angle θ relative to the equatorial plane 13 which contains the ideal geostationary orbit 14. The orbit 11 of the satellite intersects the equatorial plane 13 at two points 15 and 16, commonly called orbital nodes. The six orbital parameters used to qualify the position of a satellite are also known: the semi-major axis, the eccentricity, the inclination, the argument of the ascending node, the argument of perigee and the true anomaly. Orbit control entails quantifying these orbital parameters and carrying out the necessary operations by means of the onboard propulsion systems, to hold the satellite within a predefined area around an ideal position. As an example, for a geostationary satellite, a drift window of plus or minus 0.1°, representing a width of nearly 150 km, is assigned around a target position.
A current architecture of a satellite 10, as represented in FIG. 2, comprises a parallelepipedal structure 20 on which are fastened various devices useful to the piloting of the satellite 10 and to its mission. Telecommunications instruments 21 are installed on one face 22 whose orientation is kept facing the earth, commonly called earth face. On an opposite face 23, commonly called anti-earth face, the main satellite thruster PSP is positioned notably ensuring the thrust needed for the transfer from the low orbit to the final orbit. On two opposing lateral faces 24 and 25, commonly called north face and south face, because of their orientation relative to the equatorial plane, two sets of solar panels 26 and 27 are positioned, supplying electrical energy to the onboard systems. Various devices can be installed on the lateral faces 28 and 29, commonly called east face and west face for their orientation relative to a terrestrial longitude. Maintaining a constant orientation of the satellite relative to the earth is necessary for the mission of the satellite to run correctly, for example for the orientation of the solar panels 26 and 27 or the pointing of the telecommunication systems 21 towards the earth. This is handled by means of an attitude control system. A number of attitude control systems, suitable for detecting and correcting orientation errors, are known. Thus, the orientation of the satellite can be measured by means of a sensor assembly, comprising, for example, a sensor directed towards the earth, positioned facing the earth for a measurement on two axes, pitch and roll, relative to the earth and a set 30 of gyroscopes for detecting rotation speeds on three axes. From these measurements, corrections to the orientation of the satellite about its centre of gravity can be made, for example by means of inertia wheels 31 or of gyroscopic actuators.
A satellite equipped with such a system for controlling attitude is said to be stabilized on three axes. Typically, by controlling the speed of rotation and the orientation of the inertia wheels, an orientation error can be corrected within a reference trihedron linked to the satellite. Hereinafter, Z will be used to denote an axis directed towards the earth, also called yaw axis, Y will be used to denote an axis perpendicular to the orbit and oriented in the direction opposite to the kinetic moment of the orbit (towards the south for a geostationary), also called pitch axis, and X will be used to denote an axis forming, with Y and Z, a direct orthogonal reference frame, also called roll axis, which is oriented according to the velocity in the case of circular orbits.
For the orbit control, a number of thrusters are arranged on the structure 20 of the satellite 10. A first high-power thruster PSP, for ensuring a transfer between the initial earth orbit (after release from the launch vehicle) and the final orbit, is positioned on the anti-earth face 23. According to the known state of the art, a first set of thrusters, comprising, for example, two thrusters 32 and 33 positioned on the north face and on the south face in proximity to the anti-earth face, is used to control the inclination. A second set of thrusters, such as, for example, the thrusters 34 and 35, positioned on the east and west faces, is used to control the eccentricity and the drift. It is also known that the control of the inclination requires around five to ten times more propellant than controlling the eccentricity and the drift. For this reason, the inclination control is generally performed by means of plasma thrusters, which consume less propellant, whereas the thrusters dedicated to controlling the eccentricity and the drift are more commonly chemical ergol thrusters.
As an example, a current satellite of dry weight 2500 kg and supporting a payload of 900 kg comprises a main thruster, two plasma thrusters for the inclination and the eccentricity, and four ergol thrusters for the eccentricity and the drift. Typically, 1700 kg of ergol are needed for the initial orbit transfer, and 220 kg of Xenon are needed to ensure the orbit control of the satellite for a mission duration of approximately 15 years. Thus, the cost and the weight of the current propulsion systems limit the capacity to install a high payload. It should also be noted that, in most of the known propulsion systems for orbit control, the various thrusters installed in fact comprise two propulsive motors positioned side-by-side, for mission safety and reliability reasons. This redundancy, well known to the person skilled in the art, is not represented in the figures but it is assumed hereinafter that a thruster may be made up of one or more propulsive engines forming a propulsive assembly, and of which the thrust which can be delivered is identical, in orientation or in intensity.
FIGS. 3a, 3b and 3c illustrate the principle of orbit control for a satellite according to the known state of the art. The structure 20 of the satellite 10 is represented in a side view, the east face being visible. The thruster 32 is linked to the north face of the structure 20 by means of a two-axis mechanism 40. The two-axis mechanism 40 allows for the rotation of the thruster 32 relative to the structure 20 on a first axis parallel to the axis Y and a second axis parallel to the axis X. In FIGS. 3a to 3c, the two-axis mechanism 40 is a cardan link produced by means of a first pivot link 41 with an axis parallel to the axis Y and a second pivot link 42 with an axis parallel to the axis X. The barycentre of the satellite, situated inside the parallelepipedal structure 20, is referenced CM.
In FIG. 3a, the orientation of the thruster 32 makes it possible to direct the thrust of the thruster towards the barycentre CM of the satellite. To perform an inclination correction manoeuvre, a technique known to the person skilled in the art consists in firing the thruster 32 a first time close to an orbital node, for example 15, and then the thruster on the opposite side a second time close to the opposite orbital node, 16 in the example. Thus, the thrust from the first firing of the thruster 32, oriented towards the barycentre CM, moves the satellite in a direction that has a Z component and a Y component. Twelve hours later, the thrust from the second firing at the opposite orbital node moves the satellite in a direction that has a Z component opposite to the first firing, and which compensates the unwanted effect on the eccentricity and a Y component, also opposite, but the wanted effects of which accumulate inclination-wise. Thus, two firings of equal intensities performed twelve hours apart close to the orbital nodes 15 and 16 make it possible to cancel the effect of the radial component so as to retain only a north-south correction. This known procedure allows for a daily correction of the inclination.
By the same technique, it is also possible, by applying a second thrust of different intensity from the first, to apply eccentricity corrections on an axis perpendicular to the line joining the two orbital nodes 15 and 16. Techniques have also developed to allow for eccentricity corrections on a second axis, by offsetting the firing of the thruster relative to the orbital node, but at the cost of less effective control of the inclination. To sum up, the known systems make it possible, by means of two thruster systems 32 and 33, to ensure the control of the inclination and the control of the eccentricity on an axis with no loss of optimisation of the inclination control, or to ensure the control of the inclination and the control of the eccentricity on two axes with a loss of optimisation of the inclination control. The drift cannot be controlled by these two thrusters. A current satellite comprises, for this, four chemical ergol nozzles positioned on the east and west faces of the satellite.
The thruster systems 32 and 33 are also used to manage the quantity of movement of the attitude control systems, as illustrated in FIGS. 3b and 3c. By applying a thrust outside of the barycentre CM—in a plane Y-Z in FIG. 3b and outside of the plane Y-Z in FIG. 3c, a rotational torque is generated on the satellite—a roll torque in FIG. 3b and a pitch and yaw torque in FIG. 3c. These two torques can be used to charge or discharge the inertia wheels on two axes. For example, when the speed of rotation of an inertia wheel reaches its limit speed, efforts will be made to orient the thrust deliberately away from the barycentre CM so as to generate, in addition to the desired movement of the satellite, a torque that makes it possible to desaturate the inertia wheel, or more generally, the problem will be anticipated by adjusting the kinetic moment to desired values at the time of each manoeuvre, these desired values obviously being able to be zero, but also a value shrewdly defined so as to anticipate the trend of the kinetic moment between two manoeuvres notably under the effect of the solar radiation pressure.
It should also be noted that the barycentre of the satellite varies during the life of the satellite, notably because of the gradual consumption of the onboard propellant. In the known systems, algorithms are implemented for the combined management of the attitude control and of the orbit control, and to make it possible to take into account the position of the barycentre CM throughout the life of the satellite.
The issue of being able to have effective propulsion systems will therefore be understood. The current solutions, which implement thrusters of different types at various points on the satellite are relatively complex and costly and are very heavy, which limits the payload capacity of the satellite.